Automatic control system



Sept. 2 3, 1952' F. w. MEREDITH AUTOMATIC CONTROL. SYSTEM 5 Sheets-Sheet 1 Filed April -l, 1946 S 3552.... ofzov :Pri Qh/ Sept 23 T952 F. w. MEREDITH 2,611,559

` AUTOMATIC CONTROL SYSTEM Filed April l', 1946 3 Sheets-Sheet 5 Patented Sept. 23,

AUTOMATIC CONTROL SYSTEM Frederick William Meredith, London, England, assignor to S. Smith & Sons (England) Limited, London, England, a British company Application April 1, 1946, Serial No. 658,614 In Great Britain March 9, 1945 7 Claims. l

This invention relates to automatic control systems for maintaining a predetermined condition in which deviation of the condition from a datum value is detected and a servo-motor operated in accordance with the detected deviation to restore the condition to the datum value. In such systems provision is usually made for enabling manual control to be effected and itmay be desired that when the automatic control is put in operation the condition which is maintained is that in which it is left by the manual control. Thus in the case of an aircraft it may be desired that the automatic pilot should fly the aircraft automatically on whatever course or in whatever attitude in pitch it may have been placed manually. The object of the present invention is to provide simple and eiective means for ensuring that the datum value is that existing when the automatic control is put in operation.

In this specification and appended claims, the term manual control means the control of the condition by human means without the assistance of the automatic control system. In the case of the control of an aircraft, the expression means the control of the aircraft by the human pilot by actuation of the control surfaces without lthe assistance o f the automatic pilot and does not mean the control of the aircraft by the human pilot by manipulation of the automatic pilot.

An automatic control system for maintaining a condition ata datum value according to the present finvention comprises a device for detecting deviation of the condition from a datum value, Aa servo-motor controlled by the said device to operate means to restore the condition to the datum value, means for setting the datum value, a -switching device for putting the automatic control into and out of operation, and al second servo-motor controlled in accordance with deviations of the condition to operate the datum setting means so that the datum value changes with change of condition, the second servo-motor being rendered operative by operation of the switching device to put the automatic control out of operation whereby the datum is set to the existing condition when the automatic control is next put in operation.

``The datum-setting means may comprise the rotor of a receiver of the type known under the trade name .Selsynf the voltage distribution in the stator'of which variesv with the condition and the `said -Selsyn receiver or one coupled thereto.` vmay' lcompris'e the s'econd servo-motor and may function in the manner described in patent application Serial No. 715,707 filed December 12, 1946, now abandoned, so that the voltage induced in the rotor winding when stationary controls the first servo-motor. Such an arrangement is particularly suited for controlling the course of an aircraft, the said rotor being free to rotate during manual control and being fixed during automatic control.

Alternatively. where the condition being controlled is the angular position of a body and the device for detecting deviation is mounted on the body, the second servo-motor may be arranged to rotate the device relatively to the body about the axis of rotation of the body so that the datum value changes with the angular position of the body. Thus an aircraft may be automatically controlled in pitch by signals derived from a pendulum mounted on a platform tiltable in pitch relatively to the aircraft by the second servo-motor so that the platform is stabilised in pitch during manual control.

The invention will now be described by Way of example with reference to the accompanying drawings in which:

Figure 1 illustrates diagrammatically the layout of an automatic control system for controlling an aircraft about all three axes.

Figure 2 illustrates in further detail the rateof-turn instruments shown in Figure 1,

Figure 3 illustrates in further detail the pendulums shown in Figure 1,

Figure 4 illustrates in further detail the tangent potentiometer shown in Figure l, and

Figure 5 illustrates in further detail the control circuits shown in Figure 1.

As shown in Figure 1, the aircraft is stabilised about all three axes by means of three rate-ofturn gyroscopes. A rate-of-turn gyroscope I controls the rudder 2 to stabilise the aircraft in yaw, a compass-controlled directional gyroscope 3 serving as a course monitor during straight flight. A rate-of-turn gyroscope 5 controls the ailerons 6 to stabilise the aircraft in roll, a bank pendulum 'I serving as a bank monitor. A rate-of-turn gyroscope 8 controls the elevators 9 to stabilise the aircraft in pitch, a pitch pendulum I0 serving as a pitch monitor.

The three rate-of-turn gyroscopes I, 5 and 8 and the pitch pendulum I0 are mounted on arplatform II gimballed about the pitch axis I2 in a gimbal ring I3 pivoted in the aircraft about theroll axis I4. The platform I I can be rotated by means of a motor I5 mounted on the ring f 3 I3. The ring I3 can be rotated relatively to the aircraft about the roll aids It by a motorv I6 mounted on the aircraft. The platform II can thus be moved to any desired position in pitch or roll relatively to the aircraft by means of the two motors I5 and IS and during straight iiight conditions, the platform is centrally positioned about theroll axis.

Each of the control surfaces, rudder 2, ailerons 6 and elevators 9, is coupled through a clutch Il and gear box I8 to an electric hysteresis motor I9 of the kind described in British VPatents Nos. 576,248 and 576,249. Each ymotor I9 is Wound It will thus be seen that the control of the ailerons is similar to that of the elevators, the

for two-phase operation, one phaseia reference i phase) being coupled to an A. C. source 29, and the other phase, Vthe control phase, beingl coupled to the output of an ampliiier 2|., VEach motor I9 is coupled to an asynchronous tachometric signal generator 22 of any knownrkind. Each generator is also Wound for two-phase operation, one phase being connected to the source 23gso thatan output proportional to the speed ofthe generator and therefore of the motor is generated in the other phase.

i The rate gyroscopes will be described in greater detail hereinafter, and it Willsuiiice for an understanding ofrFigure 1 to state that a voltageis produced across the output terminals of each rate-of-turn gyroscope which is proportional to the rate-of-turn and is of the same frequency as that ofthe source 2li but is inquadrature therewith lagging or leading according as the turn is in one direction or the other. The pendulums and their electric circuits will `also, be described in greater detail hereinafter and it will sumce for an understanding of Figure 1 to state that a voltage is produced across their electrical output terminals proportional to the deviation of 'the pendulum from its central position and is of the same frequency as that of the source 29 but-is in quadrature therewith lagging or leading according as the deviation is in one direction or the other. With this preliminary explanation the operation of the individual control surfaces will noW be more fully considered.

Considering first the operation of. the elevator `thebank pendulum l.

4 rapidly stabilising any disturbance in pitch. If after the disturbance is stabilised the platform I I is not level'in pitch, the electrical signal from the pitch pendulum I0 will unbalance the system until the platform I I is level.

The operation of the aileron control will first be yconsidered under straight` flight conditions. As will be explained hereinafter, under straight night conditions a pair of contacts 23. shown `as, short-circuiting the bank pendulum l, is open.

aircraft being stabilised in roll by operating the roll as measured Yby the rate-of-turn gyroscope 5, and thev aircraft being monitored to a level positionV after vani/disturbance by signals from In this case of aileron control, however, the pendulum 1 is mounted on the aircraft and notfon the platform II.

Before considering the operation of the rudder control circuit it is necessary to say a Wordabout the components in this circuit. ns1-in the other circuits there is a rateo fturn gyroscope I measuring rate of yav v. i Mounted on, the axis I4 is a tangent potentiometer 24. This Will be, d5?- scribed in greaterd detail hereinafter and it vvill sufiice for an understandingrof vFigure 1to.state that a voltage appears across` the output terminals of the potentiometer@ Whichvlisproportional to the tangent of the angle through which the platformV i I is turned in: roll fromv its nprmal central position byl the-motor I6Y and Y iis-ofA the same frequencyV as that ofthe vsource v2l) butin quadrature therewith lagging or vleading according as the turn isv in one direction orthe other. A pendulum 25 in all respects similar to the penduiumdY is provided to detect-side slip. A signal proportional to the deviationin yaw" is obtained from the directional gyroscope 3 by means of an' electric induction transmitter-26 control it will be seen that the sum of three volt- I ages is applied to the input of the ampliiieri, of theelevator control. .These voltages are proportional respectively. to the rate of pitchof the aircraft as measured by the rate-of-turn gyroscope S, the deviation in pitch ofthe platform II from the apparent horizontal as measured by thepitch pendulum I9 and the speed ofthe motor I9. The pendulum Ill ispurely a monitor, that is it gives a comparatively Weak signal compared with those derivedfrom the rate-ofturn gyroscope 8 and the generator 22, and its effect on `the operation of the motor I9 during a disturbance may be neglected. Addition of the `voltages is effected by connecting the output of the devices 6, It, 22 in series. The motorV i9, drives the generator 22 Which generates a Yvoltageopposing the voltage derived from the device 8i. vThe motor I9 thus operates to reduce the input to the amplifier and since the gain of the amplifier Vis made very large it Vwill reduce this inputv substantially to nothing.' That is to say during a disturbance the motor will run-at such a speed that the output of the generator 22 is V,substantially equalrbut opposite to the output of the rate-of-turn device 8, or in other Words the motor I9 and therefore the elevators -9 will be operated at a speed proportional to the rate ofpitch This Will have the effect of very c and ,receiver 21 ofthe type known'underythe trade naine Selsyn. YThe rotor VWinding 'of the transmitter 25 is supplied with single-phase A; C. from a source 43 inquadrature with but ofthe same frequency as the source 20 and said; rotor is Vrotated by the directionalgyroscope 3,. lIf Vthe rotor Winding of the receiver 2l isshort-circuited, it Will follow the movementsof the transmitter rotor, and if at any instant thereceiveris `fixed and the short circuit'removed from itswinding a voltage will appear-across this lWinding proportional to the deviations of ,the transmitter rotor andv therefore of the directionalfgyroscope 3 .relatively to the 'craft fromV the positions they occupied at the instantin questions-In a similar manner signalsv proportional -to the deviations of the rudder are obtained by means of a transmitter 28 and receiver 2 9. VAs beforepshort-circuiting of the rotor Winding ofthe receiver 2l)v causes it to follow the movements vof therotor of transmitter 28 and therefore of the rudder; and xing of the rotor of receiver'29 and the removal of the short circuit causes a voltage't'o appear across `the rotor winding of the receiver sproportional to the deviations of the rudder from the position it'occupied alt the Ainstant Whenk that rotor was fixed. l

The operation of the rudder 2'wi1l'first 'be considered underl straight -fiight fconditions. Under these conditions a vpair ofr relays 30 and. 3I, shown in Figure lain theirenergised posi-- tions,v are de-energised. Whenv thexrelayiBUf is ole-energised ay brals member -32'l moves to they Y. left under the influence: of; its spring-mounting' 5.a. 30' in the drawing releasing the rotor 29 of receiver 29 and fixing the rotor 21 of receiver 21. At the same time the rotorl Winding ofv receiver 29 is short-circuited by the closing ofA switch 32a and that of receiver 21 put into circuit by the opening of switch 32h. The rotor of receiver 29 thus proceeds to follow the rudder movements as explained aboveso that its winding is maintained in a position of zero output. Meanwhile a voltage appears across the rotor winding of receiver 21 which is proportional to the deviations in yaw of the aircraft from the course it was on at the instant the relay 30 was de-,ener-` gized. When the relay 3I is de-energised the previously-mentioned contacts 23 are opened and a pair of contacts 33 closed to short circuit the electrical output of the pendulum 5.- Thus the only signals appearing in the rudder control circuit are a signal from the rate-of-turn4 eyroscope I proportional to rate of yaw, a signalproportional to deviation fromcourse, and a signal from the generator 22 proportional to rudder speed. The aeroplane is thus stabilisedagainst disturbances in yaw and is monitored to the correct course in a manner which will be obvious from the description already given of the elevator and aileron controls.

Change of course is effected by bankedturns. To achieve this the gimbal ring I3 is rotated by the motor IS and thereby the relays 30 and 3I are operated. The aircraft is thus rotated in roll at the saine rate as but in the opposite direction to the gimbal ring I3 to maintain theplatform II horizontal. When the aircraft has rotated through the desired angle of bank the motor I6 is stopped. Any disturbance of thependulum 1 during the turn 'cannot affect the aileron control as the output of .this pendulum is now short-circuted by closing of the contacts 23.

Rotation of the gimbal ring I3 will produce a signal from the tangent potentiometer 24 proportionalk to the tangent of the angle of bank ofthe aircraft. kAs is well known, for av correctly banked turn the rate of turn is directly proportional to the tangent of the angle of bank and inversely proportional to the air speed. The constant of proportionality of the tangent potentiometer 24 is selected to give a signalequal but opposite to that given by the rate-of-turn instrument I when the aircraft is turning at the correct rate and travelling at its cruising speed. In operation the aircraft will turn so as to maintain these two signals equal and will therefore turn at the speed set by the tangent potentiometer 24. Energisation of the relay 39 releases the rotor of the receiver 21 and at the same time short-circuits its winding. The rotor thus follows the movements of the directional gyroscope 3 during the turn. Energisation of relay 3I opens contacts 33 and puts the pendulum 25 in circuit to detect side slip. If the aircraft is turning at the correct speed there will be no signal from the pendulum 25. On the other hand if the aircraft is not turning at the correct speed, for example due to the air speed differing from the cruising speed, then side slip will take place and the pendulum 25 will give a signal to increase or decrease the rate of turn accordingly. The signal from the pendulum 25 detracts from the inherent damping of the system, and it is therefore necessary to introduce clamping by supplying a signal proportional to rudder displace-r ment. Thisis obtained from the rotorwinding ofthe vreceiver@ which isnow in circuit, .with

its rotor held by the member 32.

When the aircraft has turned on to the .newAVA varied by rotating the platform II in pitch by'.

means of the motor I5. As will be obvious from the description already given the aircraft will turn in pitch to maintain the platform II level. A suitable rate-of-turn gyroscope for use as gyroscope I, 5 or 8 is illustrated diagrammatically in Figure 2. As there shown a gyroscope comprises a rotor 34 mounted on spin axis 35 in a gimbalring 36 gimballed about an axis 31 on a fixed base 38. Carried by the gimbal ring 36 is a potentiometer contact39 arranged to sweep across a fixed arcuate potentiometer resistance 40 as the ring rotates relatively to the base about the axis 31. The ring 36 is restrained by a spring 4I to a central position in which the contact 39 engages the mid pointl 42 of the resistance 40. The resistance 49 is connected across the A. C. source 43. The electrical output of the gyroscope appears between the contact 39 and the mid point 42 and is in phase with or in anti-phase with source 43 in accordance with the direction of the displacement of contact 39 from the midpoint.

vIn operation any turn about an axis normal to the base 38 results in precession of the gyroscope about thev axis 31 against the spring restraint to Van extent determined by the rate of turn. Hence the output voltage will be proportional to the rate of turn. Since source 43 is in quadrature with source 20, this voltage will bef the rate of turn is in one direction or the other.v

It will be appreciated that the rate-of-turnA gyroscope described above may be arranged toV measure the rate-of-turn about any-of the three aircraft axes and that the three gyroscopes may readily be made interchangeable.

. A pendulum suitable'for use as pendulum 1, I0 or 25 is illustrated diagrammatically in Fig ure 3. As there shown a pendulum 44 of suitable conducting material is pivoted to swing-about an axis parallel to the roll axis in the case of pendulums 1 and 25 and parallel to the pitch axis in the case of pendulum I0. The pendulum 44 sweeps across an arcuate resistance 45 connected' and sweeps over a linear resistance 48 connected" across the A.-C, source 43. The output appearingvbetween thearm 41 and the mid point 49 of the resistance is thus proportional to the. gimbal tangent of the angle through which the ring I3 isturned.

In the arrangement shown in Figure 1 control is effected by means of a turn control lever 50 and a pitch control lever 5 I, the lever 53 being rotated in. roll to right orrleft to effect. turns.. to.;

7 right or left and the lever 5| being :rotated in pitch forwards or backwards to move the nose of the aircraft up or down. These two levers may conveniently be combined in the form of a'miniature ijoy stick, but their operation will be more readily understood by considering them as two separate levers. Their operation will be clear froma consideration of Figure 5. As'there shown the lever 50 is pivoted at 8l and carries a contact strip 52 insulated from the lever Y5I] vand sweeping over an arcuate potentiometer resistance 53 and a rcontact strip 54 also insulated from the lever 58 and in permanent engagement with a contact segment 55 and adapted to engage either afcontact segment V5t or a contact segment 51 accordinga's the lever is moved in one direction or the other from-its central position. Themotor IG rotates through gearing the gimbal vring i3 which 'drives a contact arm 58 Aover a potentiometer resistance 59. The motor IB is a D. C. motor and has a field winding 60 connected between th'e contact strip 52 and the contact arm 58.

The lever 5| is pivoted at 6| and carries a con# tact strip 62 sweeping over an arcuate potentiometer resistance 63. The motor llwhich rotates the platform 4 l is a D. C. motor with a iield winding 64 permanently connected across a battery 65 which supplies the energy for the control circuit. A change-over switch 66 is provided for puttingr the automatic control into and out of operation. Thisswitch has an operating coil 61 controlled by push buttons 58 and E'S'forenergising and de-energising the switch respectively. The switch is shown in the de-energised vposition, that is in the manual control position. In this position one brush of the motor l5'is connected to the pendulum through contacts 'Hl and leadll. The lead H is connected to a contact on the pendulum arm which is adapted upon displacement from its central position in one direction or the other to engage one or other 'of two segments connected respectively by leads l2 and '73 to the two terminals of the battery. The contact and segments are not shown in Figure 5 but the manner of providing them Ywill be obvious. The other brushof the motor l 5 is permanently connected to the mid point of the resistance S3. Hence any displacement of thependulum causes the motor to run in one direction or the other. The motor runs to rotate the platform ll to restorev the pendulum displacement. vHence during manual control the platform will follow the movements of the pendulum l; yConsequently when the Vautomatic control' is putl in operation itv will continue to ily, the air-` craft in the same attitude in pitch.

To put the automatic control in operation push button-68 is-pressed. This completes a circuit forv the c'oil 6l. When the switch operates, this circuit is completed through contacts T4 and the release push button .69; The automatic control can thus be instantly 'cut' out of operation by pressing the 'push button r69. For convenience ofi operation this push button maybe mounted on top Vofthe miniature joystick already mentioned.

' When'the switch 65 operates, a pairof contacts 'l5` operate'the three clutches I1 and so put the automatic control in operation. At the same time the contacts 'F0 are opened and contacts 'i6 closed.

The motor l is thus disconnected from the'pen- Consequently when the lever 5I- is' lnilts-icentral. position .'(itiisonl'y `showny displaced v to Venable the connections vto .be `clearly seen) the motor armature' is short-circuited and the motor is stationary. Movement of the lever 5| in one direction or vthe other will cause the motor to operate in one direction or the other at a speed proportionalto the displacement. Hence the aircraft can be made to change its attitudeinpitch at any desired rate while under automatic control. i

During straight ight under automatic control the lever 56 is in its central position and therelays 30 and 3l cle-energized. To change course, the ll'ever 5S is rotated in one direction or 'the other. As soon as the contact strip 512 engages the segment 56 or 51 the relays 3G and Si operate. It Will be seen that the .resistances 53 and E59 form the arms of a Wheatstone bridgeof which the field winding 4Gil forms one diagonal. ment of the lever 59- unbalances the bridge and excites the Afield so that the motor operates to restore 'the balance. Consequently the motor I6 and therefore the'gimbal ring VI3 are displaced by the', same'amount as the 'lever 5S. This determines the'angle of bank and therefore the rateof-turn so that the aircraft can be turned at any desired rateV by displacing the lever v50 by the appropriete amount.

I claim:

l. An Aautomatic system forY controlling the course of an aircraft having a rudder for control in yaw, said system comprising a rate of turn device, a servo-motor for operating said rudder, means for controlling said servo-motor so asr to operate therudder in accordance with the rate of turn measured by said device, to maintain the aircraft on 'a datum course, a heading direction sensitive device, a VElelsyn receiver the voltage dis'- tribution in thestator windings of which varies with variations of theV headingA ofthe aircraft, meansrfor controlling said servo-motor in accordance also 'with the voltage induced in the rotor winding ofsaid receiver, means for locking the rotor and means for short-circuiting the rotor, said locking 'and short-circuiting means 7oeingre-v spectivelyV operated to lock the rotor and remove the short'circuit from the rotorwhen the aircraft is'to be maintained on the datunr course determined by the position of the rotor and to unlock the rotor and applyVV the short circuit when the aircraft is to be manually controlled. r

2.v An `automatic control system for. aircraft having a rudder for control in yaw and ailerons` for controlin roll, said system comprising a rate-V of-yaw device and a rate-of-roll device mounted on a platform pivoted on the aircraft about the roll axis and tiltable at will about said roll axis, means for controlling the ailerons from said rate` of-roll deviceto stabilize the aircraft in roll. a servo-motor operating the rudder andcontrolled in accordance with the rate-of-yaw detected by said rate-of-yaw device to stabilize the aircraft 1n yaw, atangent potentiometer operated by the banking movement of the platform and producing a signal proportional to the tangent of the angle of bank of the platform, means for con-- trolling said servo-motor. in accordance also with saidsignal to effect a change of course, a heading direction sensitive device, a Selsyn receiver, the voltage distribution in thestator windings of which varies with variations of the heading of the.V aircraft, means for controlling saidA servo-motor in accordance also with the voltage induced vin the rotor windings ofV said receiver,v means Yfor locking'therotor when the. aircraft is to be main'- tainedpna datum cours'edetermined'by the position of said rotor and for short-circuiting the rotor winding and for allowing it to rotate freely and follow variations of the heading of the aircraft when the said platform is tilted in bank to effect a change of course.

3. An automatic control system for aircraft as claimed in claim 2 comprising also a second selsyn transmitter, the rotor of which is moved in accordance with movements of the rudder, a second selsyn receiver the terminals of which are connected to the corresponding-terminals of said secand transmitter, and means operative during a change of course for applying tc said servo-noter through an amplifier the signal induced in the rotor winding of said second receiver.

4. An automatic control system as claimed in claim 3 comprising also a side-slip pendulum pivoted about an axis parallel to the roll axis or" the aircraft producing a signal proportional to the lateral acceleration of the aircraft, and means for feeding the side-slip signal to said aniplier to adjust the rudder to compensate for side-slip.

5. An automatic control system for aircraft having a rudder for control in yaW and ailerons for control in roll, said system comprising a rateof-yaW device and a rate-of-roll device mounted on a platform pivoted on the aircraft about the roll axis and tiltable at will about said roll axis, means for controlling the ailerons from said rateof-roll device to stabilize the aircraft in roll, a servo-motor operating the rudder and controlled in accordance with the rate-of-yaw detected by `said rate-of-yaw device to stabilize the aircraft in yaW, a tangent potentiometer operated by the banking movement of the platform and producing a signal proportional to the tangent of the angle of bank of the platform, means for controlling said servo-motor in accordance also with said signal to eiiect a change of course, a Selsyn receiver the voltage distribution in the stator windings of which varies with variation in the heading of the aircraft, a second Selsyn receiver the voltage distribution in the stator windings of which varies with rudder movements, an amplifier to the input of which are applied the voltages generated in the rotor windings of said receivers, means for controlling said servo-motor in accordance also with the output of said amplifier and means `for locking the rotor of the rst Selsyn receiver and shorting the rotor winding of the second Selsyn receiver when the aircraft is to be maintained on a datum course determined by the settingr of the rotor of the rst receiver and for locking the rotor of the second Selsyn receiver and shorting the rotor windings of the first Selsyn receiver when a change of course is being effected.

6. An automatic control system for maintaining 9, condition at a datum value comprising a device for detecting deviation of the condition from a datum value, a Selsyn transmitter the voltage distribution in whose stator is varied in accordance with variation of the condition, a Selsyn receiver whose stator winding terminals are connected to the corresponding terminals of the transmitter windings, the voltage distribution in the Selsyn receiver stator thereby varying with the condition, a servo-motor operating means for restoring the condition to the datum value and controlled by the voltage generated in the rotor winding of said receiver and means for locking said receiver rotor in the position corresponding to the datum value at which the condition is to be maintained and means for short-circuiting the said rotor and allowing it to rotate freely and follow the variations of the condition when the system is not required to maintain the condition at its datum value, whereby the datum condition is that existing when the lock is applied and the short circuit removed.

7. An automatic control system for controlling the course of an aircraft comprising a device for detecting deviation of the aircraft in yaw from a datum course, a Selsyn transmitter the voltage distribution in whose stator is varied in accordance with variation in the course of the aircraft, a Selsyn receiver whose stator winding terminals are connected to the corresponding terminals oi the transmitter windings, the voltage distribution in the Selsyn receiver stator thereby varying with the course of the aircraft, a servo motor operating the rudder of the aircraft and controlled by the voltage generated in the rotor winding of said receiver, and means for locking said receiver rotor in the position corresponding to the datum value at which the condition is to be maintained and means for short-circuiting the said rotor and alw lowing it to rotate freely and follow the variations in the course of the aircraft when the system is not required to maintain the aircraft on a datum course, whereby the datum course is that upon which the aircraft is flying when the loci` is next applied and the short circuit removed.

FREDERICK tVILLIAl/ MEREDITH.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS Number 

